The present invention relates to fuel-propelled vehicles and, more particularly, to internal methods and arrangements for tailoring rocket exhaust plume signatures of fuel-propelled vehicles.
Fuel-propelled vehicles, such as rockets and missiles utilize rocket motors to propel the vehicle through air and space. The rocket motors general fall into three types, which are solid propellant motors, liquid propellant motors and hybrid propellant motors. Solid propellant motors utilize a solid fuel element or grain that is placed in a large solid combustion chamber. The solid fuel element or grain is usually bonded to the combustion chamber walls and burns away during flight. The liquid propellant motors employ liquid fuel tanks coupled to a fixed combustion chamber through one or more fuel lines. A hybrid propellant motor generally uses a fluid reactant (e.g., an oxidizer) to burn a solid fuel element or a fluid fuel element with a solid reactant, which are ignited in a combustion chamber.
Typically, the combustion chamber is connected to a nozzle assembly regardless of the type of rocket motor being employed. The purpose of the nozzle is to provide thrust to the vehicle by accelerating the mass of the propellants (fuel and oxidizer). The nozzle can be a supersonic nozzle with a subsonic portion and a supersonic portion. The subsonic portion is connected to the combustion chamber, while the supersonic portion opens to the outside environment. The propellant is ignited in the combustion chamber producing gases moving at subsonic speeds. The gas is then accelerated to supersonic speed by the subsonic portion of the nozzle which decreases in diameter as the gas passes through the nozzle. The gas then reaches supersonic speed and enters the supersonic portion of the nozzle which increases in diameter as the gas passes through the nozzle and exits into the environment.
The nozzle is formed from a liner and an overwrap structure for strength. The thickness of the liner is substantially greater (e.g., 90%) than the thickness of the overwrap structure. The liner insulates the overwrap structure from the heat during combustion, so that the nozzle can survive the heat from the exhaust plume during operation. The liner is formed from an ablative material, such that portions of the liner melt over time during the flight. The energy of melting the liner absorbs some of the heat from the nozzle. Additionally, the melted portions are exhausted carrying away some of the heat from the nozzle.
Chemical vehicle propulsion systems maximize performance by converting the chemical energy of a propellant into thermodynamic energy in the form of high temperature and high-pressure gases. For example, expanding the high temperature gases through a supersonic nozzle to atmospheric conditions generates a maximum thrust per pound of the propellants being utilized to drive the vehicle. Conventional rockets and missiles minimize the amount of mass expended internally to provide thermal protection of the motor/thrust high temperature components. Therefore, when designing a rocket or missile vehicle, the propellants are usually selected to maximize the high-energy release per pound, and the thermal protection materials are selected based on minimal weight and maximum thermal protection per pound. These types of propellants usually contain toxic materials (e.g., acidic compounds, oxides of toxic minerals) that are detrimental to the environment.
Target vehicles are utilized in testing rocket and missile defense systems. These target vehicles simulate an enemy vehicle so that the rocket and missile defense system can be tested prior to implementation into the field. The rocket and missile defense systems employ optics and/or infrared technology to track and destroy the target vehicle. The target vehicles operate in a similar manner to the enemy vehicles but do not carry any explosives. Additionally, due to environmental concerns, the propellants utilized in the target vehicles are nontoxic propellants. Therefore, the plume signature of a target vehicle is generally weaker than the plume signature for the actual vehicle that the target vehicle is simulating.
The following presents a simplified summary of the invention in order to provide a basic understanding of some aspects of the invention. This summary is not an extensive overview of the invention. It is intended neither to identify key or critical elements of the invention nor delineate the scope of the invention. Its sole purpose is to present some concepts of the invention in a simplified form as a prelude to the more detailed description that is presented later.
The present invention relates to methods and arrangements for tailoring rocket exhaust plume signatures of rocket exhaust systems. The tailoring of the rocket exhaust plume signature is accomplished by providing and locating at least one structure having materials and/or additives that modify the radiant intensity pattern of the rocket exhaust plume upon ablation or melting into the rocket exhaust plume. The materials and/or additives can be manufactured or incorporated into preexisting ablative materials of the rocket exhaust system or be provided as stand-alone structures.
In one aspect of the invention, the material or additives can be selected to provide a desired plume signature by configuring the materials in a specific manner. The materials or additives can be configured to provide ablation gradients in the ablating structure. The ablation gradients define the varying effects of the configured plume modifying materials on the radiant intensity of the exhaust plume. Therefore, plume signature tailoring can be achieved by structurally varying materials and/or compositions of different materials or additives within the ablating structure to modify (e.g., enhance, reduce, resonate) the radiant intensity or infrared signature associated with the exhaust plume.
In one aspect of the invention, the fuel-propelled vehicle is a self-propelled vehicle, such as a missile or rocket. The missile or rocket can be a target vehicle used in testing of missile defense systems. One or more additives can be provided in a nozzle liner that insulates a nozzle overwrap structure to protect the nozzle overwrap structure from the heat of the nozzle exhaust. The one or more additives ablate with the insulating material into the exhaust plume, modifying the radiant intensities of the exhaust plume based on the manner that the additives are configured in the liner. Alternatively, the one or more additives can be provided in jet vanes of the missile or rocket, or in stand alone structures disposed within or near the nozzle exhaust.
To the accomplishment of the foregoing and related ends, certain illustrative aspects of the invention are described herein in connection with the following description and the annexed drawings. These aspects are indicative, however, of but a few of the various ways in which the principles of the invention may be employed and the present invention is intended to include all such aspects and their equivalents. Other advantages and novel features of the invention will become apparent from the following detailed description of the invention when considered in conjunction with the drawings.